This invention relates to aircraft blades and in particular, to a new and improved apparatus for use in continuously detecting cracks in propeller and helicopter rotor blades.
Aircraft blades are subjected to severe stress and occasionally develop minute cracks. It is of critical importance that the crack in the blade be detected at an early time so that the blade may be replaced preventing an inflight accident. A wide variety of methods are available for detecting cracks, including the making of x-ray pictures and the magnetic flux techniques. However, these require considerable equipment and can only be performed when the aircraft is at rest.
Another system provides a pressure differential in a sealed hollow blade, with the interior blade pressure being either above or below atmospheric. A pressure sensor is mounted on the blade and provides a visual indication of the pressure differential, with a drop in the differential indicating leakage due to a crack. This type of device provides for blade integrity measurement without requiring removal of the blade from the aircraft. However, the visual inspection can be performed only when the aircraft is on the ground. It has been suggested that the blade internal pressure could be transmitted to an indicator in the cockpit while the aircraft is in flight, utilizing a set of slip rings at the rotating hub for information transmission. However, slip ring systems present a number of problems. They are susceptible to dirt and grease, they are difficult to install and maintain, and they increase the cost and complexity of an already complex rotor shaft system.
An improved apparatus for continuously indicating the condition of hollow aircraft blades while the aircraft is in flight is shown in U.S. Pat. No. 3,985,318. This apparatus does not require any connection between the rotating blade and the remainder of the aircraft and can be utilized to give a go-no go indication for safe and warning conditions, and/or proportional type indications as a measure of pressure differential. The device is mounted on the blade and includes a shaft which moves as a function of pressure within the blade. A radiation source is carried on the shaft and is moved between shielded and unshielded positions.
However, the radiation source is always within the housing, which reduces the sensitivity of the overall system.
In an improved version now in use, the moving member which carries the radiation source has an extension which projects above the housing, with the radiation source carried on this extension. When the instrument is in the safe or reset condition, the portion of the moving member carrying the radiation source is within the housing, which serves as a shield for the radiation. When the instrument is indicating an unsafe condition, such as a low pressure within the blade resulting from leakage through a crack, the moving member is moved outward exposing the radiation source. With this arrangement, the magnitude of the radiation source necessary for operation is less than that required for the earlier design where the radiation source was always within the housing.
A cover or shield is provided for the instrument when the aircraft is not in use. A typical cover is a sleeve closed at the upper end for sliding downward over the instrument, with the closed end of the sleeve engaging the projecting member to push on and maintain the projecting member in the safe or reset position. However problems have been encountered with this construction, with significant damage to the instrument when the cover is not carefully utilized.
Accordingly, it is an object of the present invention to provide a new and improved cover for a helicopter blade crack indicator, which cover functions as a protective cover, a radiation shield, and a test unit for the indicator.
Other objects, advantages, features and results will more fully appear in the course of the following description.